1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to film cooling holes in a turbine airfoil.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
A gas turbine engine burns a fuel to produce a hot gas flow for the production of mechanical power. Compressed air from a compressor is mixed with a fuel in the combustor to produce the hot gas flow. The hot gas flow is then passed through a turbine having multiple stages of stator vanes and rotor blades to extract mechanical power from the flow. The engine efficiency can be increased by increasing the hot gas flow into the turbine. However, the turbine parts—especially the first stage stator vanes and rotor blades—are exposed to the hottest temperature and therefore the limiting properties of these parts determine the highest temperature entering the turbine.
One method of increasing the temperature into the turbine is to provide complex cooling circuits within the airfoils (stator vanes, rotor blades). Complex multi-pass serpentine cooling flow circuits have been proposed to provide high levels of cooling using low amounts of cooling air. Since the cooling air used within the airfoils is generally bleed off air from the compressor, maximizing the cooling effect of the air while minimizing the amount of air used will also increase the engine efficiency. Besides the internal cooling circuit within an airfoil film cooling holes are also used to provide a film of cool air on the surface of the hottest parts of the airfoil. Axial and radial cooling holes have been used to provide film cooling to the airfoil.
To achieve a high cooling effectiveness, cooling air discharged through film cooling holes must be deflected as rapidly as possible and flow in a protective manner along the profile surface of the airfoil. To protect zones lying between the bores, rapid lateral spreading of the cooling air is also necessary. This is achieved by using a diffuser with the cooling holes which due to the lateral widening permits a wider area of the airfoil surface to be covered. The diffuser also lowers the velocity of the cooling air from the hole so that the cooling air does not blow out from the film layer and into the hot gas flow. Geometric diffuser forms in which the bore hole is widened not only laterally but also on the downstream side of the hole are used to further improve the mixing behavior. The blow-out rates in these diffuser holes are small so that there is little risk of the cooling air passing through the flow boundary layer.
Ideally, it is desired to bathe 100% of the airfoil surface with a film of cooling air. However, the cooling air leaving the cooling hole exit generally forms a cooling film stripe no wider than or hardly wider than the dimension of the cooling hole exit perpendicular to the hot gas flow. Limitations of the number, size and spacing of the cooling passages results in gaps in the protective film and/or areas of low film cooling effectiveness which produce localized hot spots on the airfoil. Airfoil hot spots are one factor which limits the operating temperature of the engine.
A standard film cooling hole 14 is shown in FIG. 1 which passes straight through the airfoil wall 12 at a constant diameter and exits at an angle to the surface 13. FIG. 2a shows a top view of this film cooling hole 15, and FIG. 2b shows a side view. Because the cooling hole 14 is not perpendicular to the airfoil surface 13, the opening 15 of the hole has a longer major axis than the minor axis as shown in FIG. 2a. Some of the cooling air is consequently ejected directly into the mainstream causing turbulence, coolant dilution, and loss of downstream film effectiveness. In addition, the hole breakout in the stream-wise elliptical shape will induce stress problems in the blade application. The stress field is shown in FIG. 1 as reference numeral 17 in this film cooling hole design. The film of cooling air is represented by 16 as it is discharged from the hole 15.